The present invention generally relates to attitude control arrangements for a spacecraft, and more particularly, to improvements in controlling operation of a set of rigidly attached chemical thrusters and a gimballed momentum wheel.
As a spacecraft orbits the Earth, the position of the spacecraft relative to the Earth must be periodically corrected by firing thrusters to maintain the spacecraft in the desired orbit. During an orbit correction, the spacecraft""s attitude must be maintained in a desired direction.
A spacecraft 10 equipped with attitude and orbit control devices is shown in FIG. 1. The orbit of the spacecraft 10 is controlled using thrusters while errors are measured relative to inertial space using gyroscopes. Attitude is controlled using either a gimballed momentum wheel 14 or the thrusters. The computations required to command the wheel or thrusters are performed by an onboard computer (not shown). Alternatively, the spacecraft may use a reaction wheel system (not shown) in place of the gimballed momentum wheel 14. In general, any torque actuation system which has limited control authority and momentum storage capacity on all three axes can be employed.
A spacecraft control system which uses both the thrusters and the gimballed momentum wheel is commonly called Transition Mode because the system is used in transition from a Stationkeeping Mode to a Normal Mode. Transition mode without thrusters is called Wheel Transition Mode. The Stationkeeping Mode is executed by the control system and the attitude is controlled using errors measured relative to inertial space and thrusters to both maintain the attitude of the spacecraft and change the orbital velocity. The Normal Mode is executed after Transition Mode by the control system using azimuth and elevation errors and the gimballed momentum wheel 14 to control the spacecraft attitude as it orbits the Earth with a constant velocity.
In the Normal Mode, the gimballed momentum wheel 14 controls the spacecraft attitude by changing the momentum of the spacecraft about its pitch axis 16, its roll axis 18 and its yaw axis 20. The amount of momentum on a given axis is controlled by changing the rotational speed xcfx89w of the momentum wheel 14, the roll gimbal angle xcfx81, and the yaw gimbal angle 65 . The inertia of the momentum wheel Ixcfx89; is constant. FIG. 2 shows a momentum wheel and the relationship of the momentum wheel relative to the gimbal angles xcfx81 and xcex3 and the rotational speed xcfx89w about the pitch axis 16 as defined herein. The moment xe2x80x9cHxe2x80x9d of the spacecraft (ignoring the orbit spin rate) about the roll and yaw and pitch axes for any given wheel speed xcfx89w and gimbal angles xcfx81 and xcex3 is as follows:
Hroll=Iwxcfx89w sin xcex3cos xcfx81≈Iwxcfx89wxcex3xe2x80x83xe2x80x83(1)
Hpitch=xe2x88x92Iwxcfx89w cos xcfx81cos xcex3≈xe2x88x92Iwxcfx89wxe2x80x83xe2x80x83(2)
xe2x80x83Hyaw=xe2x88x92Iwxcfx89w sin xcfx81cos xcex3≈xe2x88x92Iwxcfx89wxcfx81xe2x80x83xe2x80x83(3)
The torques, Troll, Tpitch, and Tyaw, imparted to the spacecraft 10 by the momentum wheel are given by Euler""s equation. For small attitude errors, for example, errors less than 5 degrees, the torques imparted to the spacecraft by the momentum wheel 14 are:                               T          roll                =                              Δ            ⁢                          xe2x80x83                        ⁢                          H              roll                                            Δ            ⁢                          xe2x80x83                        ⁢            t                                              (        4        )                                          T          pitch                =                              Δ            ⁢                          xe2x80x83                        ⁢                          H              pitch                                            Δ            ⁢                          xe2x80x83                        ⁢            t                                              (        5        )                                          T          yaw                =                              Δ            ⁢                          xe2x80x83                        ⁢                          H              roll                                            Δ            ⁢                          xe2x80x83                        ⁢            t                                              (        6        )            
These results are used to evaluate the momentum changes, and the torques required to manage the momentum.
A1ternatively, the torques may be applied to the spacecraft using thrusters, as shown in FIG. 1. Generally, the thrusters are chemical thrusters of conventional design to produce force levels between 2N and 50N. The thrusters are equipped with valves to switch the force on and off, and the thrusters are not gimballed, but rather rigidly attached to the spacecraft. The thrusters are throttled only to the extent that they may be switched at a duty cycle such that the average force is less than the force associated with continuous operation. Each thruster on the spacecraft applies some torque about each axis. Usually, the thrusters are placed on the spacecraft 10 so that they may be fired in pairs to impart a nearly pure torque about a selected one of the axes, or nearly a pure translation along one of the axes.
Referring back to FIG. 1, thrusters N1-N4 are called North thrusters, thrusters A1-A4 are called Axial thrusters, thrusters E1 and E2 are called East thrusters and thrusters W1 and W2 are called West thrusters. Firing the North thrusters N1-N4 at the orbital nodes reduces orbital inclination. The north thrusters may also be used to apply roll and yaw torques to control the attitude of the spacecraft. The axial thrusters A1-A4 are used to control the pitch and the roll attitude of the spacecraft but do not contribute to orbit inclination. East thrusters E1 and E2 correct for an east drift of the spacecraft and the West thrusters W1 and W2 correct for a west drift of the spacecraft. For a 3000 lb. spacecraft equipped with 5 lb. thrusters in geosynchronous orbit, the inclination maneuvers typically require 60 seconds of continuous thruster fire using 2 thrusters, while the drift maneuver require 3 seconds of continuous thruster fire using two thrusters.
Thruster attitude control has several problems. First, it is necessary to rapidly switch the thrusters on and off as attitude errors develop and subside to deliver sufficiently small torque over an extended period of time. This induces jarring accelerations on the spacecraft which may adversely effect the spacecraft""s structural or mechanical elements. Secondly, switching a thruster on and off to deliver small attitude correction torques is limited by the capability/design of the switching mechanism and propellant flow limitations. Hence, the control loop contains a control deadband. Control deadbands tend to induce limit cycle instabilities, and place extremne computational pressures on the throughput of the onboard electronics due to the computational requirements for both selecting and firing the appropriate thrusters.
In addition, the maneuver thrusters generally are not directed exactly through the center of mass of the spacecraft. Therefore, the stationkeeping maneuver adds momentum to the spacecraft. This additional momentum results in attitude errors which are removed by the control system by firing other thrusters or by modulating the maneuver thrusters. Hence, the thrusters are performing attitude control and momentum control simultaneously. The disadvantage of this is that an undesirably high number of extremely short thruster pulses are required from both the north and axial thrusters.
Lastly, the thruster control system must fire axial thrusters in excess of the needs of momentum management, thereby wasting thruster fuel. More specifically, a thruster firing control arrangement which uses axial thrusters A1-A4 (which are orthogonal to the direction of the orbital velocity change) to control the attitude of the spacecraft 10 during the stationkeeping mode wastes thruster fuel. Similarly, all extremely short thruster pulses ( less than 300 msec) are wasteful because thrusters are inefficient for such short fire durations. Unfortunately, the gimbals of the momentum wheel 14 do not have sufficient range to maintain the attitude over a continuous drift or inclination maneuver, so the wheel speed and gimbals are conventionally held constant while the thrusters are used to control the attitude during the thruster maneuvers.
Because the use of the axial thrusters and other short north thruster pulses for momentum management and attitude control of the spacecraft during the Stationkeeping Mode is wasteful of thruster fuel, a need exists for an attitude control arrangement which eliminates as much as possible the use of extremely short north and axial thruster pulses during the Stationkeeping Mode, thereby decreasing the fuel requirements and increasing the life of the spacecraft.
Therefore, it is an object of the present invention to provide a new stationkeeping method and system which eliminates the use of thrusters to maintain the spacecraft attitude during the execution of the stationkeeping mode.
A further object of the present invention is to provide a new stationkeeping method and system which separates maneuver, momentum, and attitude control operations so that a momentum wheel can perform attitude control while a set of thrusters can perform the maneuver and momentum control operations.
Another object of the present invention is to reduce overall fuel usage without increasing attitude errors or excessively increasing the duration of the overall stationkeeping process (less than a half hour).
Yet another object of the present invention is to greatly simplify the real-time operation of stationkeeping maneuvers, thereby reducing throughput requirements on a control processor.
In accordance with these and other objects, the present invention provides a xe2x80x9cTransition Mode Stationkeepingxe2x80x9d (TMS) substitute for a conventional stationkeeping mode. Under TMS, an orbiting spacecraft would have thrusters which fire to correct orbital errors and manage the total momentum, and a gimballed momentum wheel (or other momentum storage device) to maintain the attitude. A complete maneuver is divided into short duration open-loop maneuver thruster pulses which are combined with closed-loop momentum management thruster pulses. The maneuver pulses are fired open-loop by a maneuver sequencer to correct orbital errors of the spacecraft, while a closed-loop momentum management processor pulses the thrusters to manage the momentum, i.e., level the platform and maintain the rotational speed of the momentum wheel.
To avoid attitude errors in the case of a planned thruster firing not occurring, a thruster phase plane controller may be set up with sufficiently large deadbands so that it does not interfere with correct operation of TMS. The phase plane controller could be operated on the secondary processor to avoid throughput problems.
Thus, the present invention provides a spacecraft attitude control method and system which simultaneously: (1) changes the orbital velocity using fixed chemical thrusters; (2) maintains inertial attitude using momentum storage devices; and (3) manages the momentum in the momentum storage devices. The present invention is particularly applicable to satellites equipped with chemical thrusters that are rigidly attached to provide a force in a predetermined direction relative to the satellite, a gimbaled momentum wheel or other 3-axis momentum storage device for attitude control, a 3-axis inertial reference such as orthogonal gyros, and a processor to compute feedback control. Furthermore, in the preferred embodiment, the maneuver thrusters are augmented by other similar thrusters with the authority necessary to offset the momentum induced by the maneuver thrusters.
The above objects and other objects, features, and advantages of the present invention are readily apparent from the following detailed description of the best mode for carrying out the invention when taken in connection with the accompanying drawings.